1. Field of the Invention
The present invention relates to composite structures and more particularly to filament-reinforced composite panels which uniquely incorporate a particular corrugated core between filament-reinforced outer sheets.
2. Description of the Prior Art
With the advent of filament-reinforced composites, the potential of significant improvements in structural materials designed to meet the rigid requirements of space-age hardware was established. The stiffness, strength and light weight of such composites make them eminently suitable for aerospace applications where weight savings result in enormous benefits. Some early discussions of a wide variety of composites appear in Industrial Research, October 1969, pp. 58-85.
Although an intense area of interest and activity has been the consolidation and utilization of filament-reinforced materials of the type described above, some limitations have been recognized. One problem, for example, occurs when a conventional and relatively thick aircraft skin panel made of isotropic metal such as aluminum, steel, titanium, or the like is replaced by a relatively thin sheet of a filament-reinforced composite such as boron-aluminum or graphite-epoxy. Because of its extremely reduced thickness, the composite sheet has reduced buckling resistance and low panel further stiffness which renders it subject to unacceptable deflection, flutter, vibration and resonance during aircraft operation. To obviate these and other drawbacks, it has been suggested that making sandwich panels without undue weight increase is necessary. Unfortunately however common approaches to sandwich construction may produce other problems. It was found for example that the use of honeycomb as a core material results in saddling, i.e., out of plane warping, when that material is formed into a simple curved shape. Other structures, such as those taught in U.S. Pat. Nos. 3,995,081 and 4,051,289, utilize composite materials throughout and for various reasons are not considered satisfactory.